Gas turbine engines are generally used in a wide range of applications, such as aircraft engines and auxiliary power units. In a gas turbine engine, air is compressed in a compressor, mixed with fuel, and ignited in a combustor to generate hot combustion gases, which flow downstream into a turbine section. In a typical configuration, the turbine section includes airfoils, such as stator vanes and rotor blades, disposed in an alternating sequence along the axial length of a generally annular hot gas flow path. The rotor blades are mounted at the periphery of one or more rotor disks that are coupled in turn to a main engine shaft. Hot combustion gases are delivered from the engine combustor to the annular hot gas flow path, thus resulting in rotary driving of the rotor disks to provide an engine output.
Due to the high temperatures in many gas turbine engine applications, it is desirable to regulate the operating temperature of certain engine components, particularly those within the mainstream hot gas flow path in order to prevent overheating and potential mechanical issues attributable thereto. For example, it is desirable to cool the shroud in the turbine section (i.e., the turbine shroud) to prevent or reduce oxidation, thermo-mechanical fatigue, and/or other adverse impacts. However, given the high temperature of engine operation, cooling remains a challenge.
Accordingly, it is desirable to provide gas turbine engines with improved turbine shroud cooling. Furthermore, other desirable features and characteristics of the present disclosure will become apparent from the subsequent detailed description and the appended claims, taken in conjunction with the accompanying drawings and this background discussion.